The optimization of working structures makes it necessary to adapt the local section of said structures to the intensity of the flow of forces through them. Thus, in the case of a metal structure, constituting particularly a fuselage panel, a rib or a longitudinal member in the structure of an aircraft, such adaptations are achieved by material removal, so as to create thicker and more rigid areas capable of transmitting higher force flows in the areas where greater rigidity and/or mechanical strength is required. At the location of cross section variations, geometric stress concentrations are created, which are related to the discontinuity in rigidity between the thicker and the thinner areas. These geometric discontinuities are thus the main locations for the initiation of fatigue cracks. This phenomenon is taken into account by a stress concentration coefficient or Kt, which coefficient increases as the distance over which the section varies is shorter.
If this same type of part is made of a laminated composite material, with a fiber reinforcement, said part is not obtained by material removal but by laying up fibrous layers or plies. If the local reinforcement of such a part by increasing its thickness remains desirable, it is achieved by the local addition of plies or patches extending over the area to reinforce. As with metal parts, the difference in rigidity between the running section and the reinforced section leads to the geometric concentration of stresses at the section variation location. Because composite materials have few mechanisms capable of accommodating the propagation of a crack, these coefficients may even have significantly higher values than with metal materials, with a given geometry. What is more, if a patch is made up of a stack of plies, there is a critical mode of degradation where the excess thickness corresponding to said patch is simply sheared at the interface of the patch with the remainder of the part. That phenomenon, also called ‘peeling’ is all the more likely when the said interface has defects such as porosities, or more generally bonding defects. Thus, in order to remedy these unwanted effects, patches are made by applying very gradual plies that lead to connecting slopes between the surface of the part and the excess thickness. These slopes generally range between 0.02 and 0.05. Thus, a thickness variation of 1 mm is applied over a distance ranging between 20 mm and 50 mm, so that the patch extends over a large surface and the reinforcement is not very localized in terms of geometry. Further, said patches are frequently covered with a ply extending between the skin and the top of said patch. Alternatively, if greater slopes are required, the adding of fastening elements such as rivets that go through the patch and the skin are known in the prior art. The document FR-2933067-A thus describes different solutions for the localized reinforcement of a panel made of laminated composite material. These solutions of the prior art, which allow the use of tape laying for making the panels, have drawbacks in terms of mass. Even though these solutions of the prior art have been developed initially for composite materials with a thermosetting matrix, these same solutions are reproduced for composite materials with a thermoplastic matrix, because the phenomena explained above are primarily attributable to the laminated nature of the material.
The document WO 01/58680 describes a structural panel comprising reinforcement patches, where both the panel and the patch are made of a composite with fiber reinforcement in a thermosetting resin.